ONERA OA212 AIRFOIL (oa212-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: ONERA OA212 AIRFOIL (oa212-il) Reynolds number: 100,000 Max Cl/Cd: 44.3 at α=9.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-oa212-il-100000.txt Download as CSV file: xf-oa212-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: ONERA OA212 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4733 0.10602 0.10134 -0.0053 1.0000 0.1222
-9.000 -0.5054 0.10100 0.09639 -0.0193 1.0000 0.1270
-8.750 -0.5246 0.09503 0.09041 -0.0265 1.0000 0.1282
-8.500 -0.4680 0.09292 0.08833 -0.0159 1.0000 0.1322
-8.250 -0.4595 0.08947 0.08488 -0.0179 1.0000 0.1372
-8.000 -0.4869 0.08453 0.07994 -0.0275 1.0000 0.1423
-7.750 -0.5418 0.08315 0.07827 -0.0304 1.0000 0.1440
-7.500 -0.4996 0.07710 0.07253 -0.0280 1.0000 0.1473
-7.250 -0.4991 0.07515 0.07062 -0.0251 1.0000 0.1500
-7.000 -0.5110 0.07327 0.06875 -0.0224 1.0000 0.1529
-6.750 -0.5541 0.07277 0.06794 -0.0208 1.0000 0.1607
-6.500 -0.5595 0.06858 0.06376 -0.0189 1.0000 0.1629
-6.250 -0.5521 0.06572 0.06101 -0.0164 1.0000 0.1650
-6.000 -0.5486 0.06354 0.05885 -0.0140 1.0000 0.1685
-5.750 -0.5479 0.06049 0.05547 -0.0153 0.9976 0.1812
-5.500 -0.5227 0.04639 0.03988 -0.0203 0.9925 0.1109
-5.250 -0.4948 0.04204 0.03535 -0.0222 0.9879 0.1071
-5.000 -0.4636 0.03642 0.02834 -0.0233 0.9836 0.0980
-4.750 -0.4342 0.03371 0.02566 -0.0250 0.9787 0.1011
-4.500 -0.3960 0.03200 0.02372 -0.0276 0.9745 0.1050
-4.250 -0.3648 0.03020 0.02152 -0.0285 0.9688 0.1076
-4.000 -0.3254 0.02868 0.01952 -0.0307 0.9645 0.1132
-3.750 -0.2945 0.02753 0.01837 -0.0319 0.9584 0.1185
-3.500 -0.2533 0.02680 0.01733 -0.0343 0.9535 0.1270
-3.250 -0.2226 0.02574 0.01634 -0.0353 0.9470 0.1348
-3.000 -0.1803 0.02493 0.01543 -0.0381 0.9420 0.1468
-2.750 -0.1475 0.02447 0.01496 -0.0392 0.9339 0.1596
-2.500 -0.0996 0.02370 0.01425 -0.0428 0.9263 0.1784
-2.250 -0.0607 0.02294 0.01369 -0.0445 0.9161 0.1995
-2.000 -0.0322 0.02241 0.01338 -0.0444 0.9048 0.2236
-1.750 -0.0039 0.02189 0.01314 -0.0442 0.8956 0.2620
-1.500 0.0183 0.02030 0.01281 -0.0431 0.8874 0.4629
-1.250 0.2138 0.02016 0.01359 -0.0668 0.8939 1.0000
-1.000 0.2334 0.02028 0.01358 -0.0650 0.8834 1.0000
-0.750 0.2528 0.02035 0.01355 -0.0634 0.8714 1.0000
-0.500 0.2724 0.02043 0.01355 -0.0620 0.8594 1.0000
-0.250 0.2909 0.02041 0.01344 -0.0596 0.8492 1.0000
0.000 0.3101 0.02036 0.01332 -0.0577 0.8373 1.0000
0.250 0.3297 0.02033 0.01324 -0.0559 0.8250 1.0000
0.500 0.3466 0.02009 0.01291 -0.0527 0.8160 1.0000
0.750 0.3666 0.01998 0.01276 -0.0510 0.8026 1.0000
1.000 0.3859 0.01983 0.01257 -0.0489 0.7899 1.0000
1.250 0.4020 0.01939 0.01206 -0.0452 0.7814 1.0000
1.500 0.4223 0.01920 0.01185 -0.0433 0.7669 1.0000
1.750 0.4419 0.01895 0.01156 -0.0410 0.7529 1.0000
2.000 0.4608 0.01860 0.01117 -0.0383 0.7401 1.0000
2.250 0.4784 0.01810 0.01059 -0.0349 0.7289 1.0000
2.500 0.4988 0.01779 0.01024 -0.0327 0.7126 1.0000
2.750 0.5192 0.01748 0.00988 -0.0304 0.6957 1.0000
3.000 0.5399 0.01717 0.00951 -0.0281 0.6779 1.0000
3.250 0.5608 0.01690 0.00915 -0.0259 0.6593 1.0000
3.500 0.5833 0.01678 0.00896 -0.0243 0.6376 1.0000
3.750 0.6061 0.01673 0.00884 -0.0229 0.6151 1.0000
4.000 0.6289 0.01673 0.00872 -0.0215 0.5940 1.0000
4.250 0.6521 0.01682 0.00868 -0.0202 0.5742 1.0000
4.500 0.6755 0.01699 0.00873 -0.0191 0.5553 1.0000
4.750 0.6992 0.01724 0.00887 -0.0182 0.5375 1.0000
5.000 0.7231 0.01755 0.00910 -0.0174 0.5206 1.0000
5.250 0.7471 0.01791 0.00940 -0.0167 0.5047 1.0000
5.500 0.7710 0.01832 0.00979 -0.0161 0.4898 1.0000
5.750 0.7947 0.01876 0.01021 -0.0154 0.4761 1.0000
6.000 0.8183 0.01922 0.01061 -0.0147 0.4636 1.0000
6.250 0.8420 0.01968 0.01099 -0.0139 0.4519 1.0000
6.500 0.8658 0.02023 0.01160 -0.0134 0.4389 1.0000
6.750 0.8897 0.02080 0.01218 -0.0129 0.4272 1.0000
7.000 0.9140 0.02133 0.01261 -0.0122 0.4161 1.0000
7.250 0.9379 0.02190 0.01329 -0.0118 0.4029 1.0000
7.500 0.9617 0.02249 0.01394 -0.0114 0.3899 1.0000
7.750 0.9856 0.02302 0.01446 -0.0109 0.3764 1.0000
8.000 1.0093 0.02345 0.01484 -0.0104 0.3615 1.0000
8.250 1.0328 0.02384 0.01517 -0.0098 0.3459 1.0000
8.500 1.0553 0.02420 0.01559 -0.0093 0.3299 1.0000
8.750 1.0771 0.02458 0.01609 -0.0088 0.3140 1.0000
9.000 1.0982 0.02495 0.01654 -0.0082 0.2976 1.0000
9.250 1.1181 0.02530 0.01696 -0.0075 0.2804 1.0000
9.500 1.1364 0.02565 0.01737 -0.0066 0.2617 1.0000
9.750 1.1506 0.02613 0.01803 -0.0054 0.2381 1.0000
10.000 1.1607 0.02684 0.01873 -0.0039 0.2097 1.0000
10.250 1.1641 0.02816 0.01992 -0.0018 0.1751 1.0000
10.500 1.1624 0.03006 0.02160 0.0006 0.1471 1.0000
10.750 1.1603 0.03210 0.02342 0.0030 0.1292 1.0000
11.000 1.1622 0.03416 0.02535 0.0048 0.1159 1.0000
11.250 1.1680 0.03610 0.02727 0.0064 0.1057 1.0000
11.500 1.1761 0.03804 0.02919 0.0078 0.0979 1.0000
11.750 1.1881 0.03987 0.03082 0.0089 0.0914 1.0000
12.000 1.1992 0.04174 0.03290 0.0100 0.0864 1.0000
12.250 1.2157 0.04343 0.03447 0.0108 0.0818 1.0000
12.500 1.2309 0.04548 0.03662 0.0117 0.0780 1.0000
12.750 1.2406 0.04754 0.03885 0.0127 0.0749 1.0000
13.000 1.2610 0.04941 0.04069 0.0132 0.0719 1.0000
13.250 1.2806 0.05204 0.04340 0.0137 0.0697 1.0000
13.500 1.2802 0.05480 0.04648 0.0149 0.0685 1.0000
13.750 1.2787 0.05776 0.04972 0.0158 0.0671 1.0000
14.000 1.2771 0.06080 0.05298 0.0164 0.0658 1.0000
14.250 1.2781 0.06375 0.05609 0.0169 0.0645 1.0000
14.500 1.2779 0.06692 0.05941 0.0172 0.0634 1.0000
14.750 1.2775 0.07046 0.06308 0.0173 0.0625 1.0000
15.000 1.2713 0.07468 0.06749 0.0172 0.0621 1.0000
15.250 1.2559 0.07958 0.07264 0.0165 0.0619 1.0000
15.500 1.2291 0.08535 0.07872 0.0147 0.0621 1.0000
15.750 1.1682 0.09573 0.08961 0.0087 0.0634 1.0000
16.000 0.9841 0.13705 0.13169 -0.0198 0.0719 1.0000
16.250 0.9861 0.14110 0.13574 -0.0211 0.0709 1.0000
|
Polar data table (+)
Polar graphs
<< Back to ONERA OA212 AIRFOIL (oa212-il)